Method for navigating a robot and arrangement at said robot

ABSTRACT

The present invention relates to an arrangement and a method for guiding a missile ( 1 ) towards a target ( 2 ). The missile has information on its own position, velocity vector and future velocity profile and continuously receives information on the position and velocity vector of the target. The invention is characterized in that from the information on the missile and the target a point of interception (A) is predicted at which the missile is expected to strike the target. A flight time is then calculated which indicates the time that it will take for the missile to travel to the predicted point of interception (A). In addition a fictitious point of interception (B) is calculated which is situated at a higher altitude than the predicted point of interception and the distance of which to that point is related to the calculated flight time. Finally the missile velocity vector is directed towards the said fictitious point.

The present invention relates to a method of guiding a missile towards atarget according to the pre-characterising part of claim 1.

The present invention also relates to an arrangement in a missile forguiding the said missile towards a target according to thepre-characterising part of claim 4.

Several earlier methods of guiding a missile towards a target alreadyexist. One method frequently encountered is so-called visual bearingguidance. Furthermore, there are methods of supplementing the visualbearing guidance with some function for enhancing the trajectory profilein the vertical plane for the purpose of obtaining increasedperformance. By means of this function the missile aims slightly higherthan indicated by the line of sight, thereby producing a curvature ofthe missile trajectory.

U.S. Pat. No. 5,082,200 relates to a method of guiding a missile towardsa point of interception with a target in the form, for example, of asatellite. A characteristic of the target, however, is that it has aknown or predictable trajectory profile. The course of the missile ismade to deviate from a collision course with the target by an angledependent upon the distance to the target. The further the missile isfrom the target, the greater the angle of deviation.

This method is used for missiles propelled by solid fuel engines, inwhich the engines are fired at certain intervals and in which the engineconstitutes the sole control element for guiding the missile.

One object of the present invention is to produce a method and anarrangement for guiding a missile towards a target that represent animprovement on the earlier known system described above.

The object has been achieved by means of a method for guiding a missiletowards a target that has the characteristics specified in claim 1.

An arrangement for performing the said method has the characteristicsspecified in claim 4.

Preferred methods and embodiments also have any or some of thecharacteristics specified in the subordinate claim for each claimcategory.

The method and the arrangement according to the invention have variousadvantages:

The missile is guided towards a point above the target in order to makeuse of the lower air resistance at higher altitude, with a view tooptimising the trajectory of the missile in respect of any chosencriterion, for example maximum final velocity or maximum averagevelocity, or for the lowest possible fuel consumption. The latter may beutilised, among other things, if the missile calculates that it cannotreach the target by flying a straight trajectory, since the missile canclimb to an altitude at which the fuel consumption is lower and in thisway increase the range. Alternatively the missile maintains thataltitude in order, simply via an active homing device, to activate thetarget's missile warning system for as long a time as possible.

The missile can be manoeuvred on to a trajectory such that the kinematicperformance of the missile is improved compared to previously knownmethods.

As stated earlier, the method and the arrangement offer the freedom toselect the optimisation criterion and final velocity conditions. Inaddition the trajectory of the missile can be adjusted as the scenariodevelops (interception point migration).

The method and arrangement according to the invention permit speculationon target behaviour, such speculation being used in estimating the pointof interception. In this way the target behaviour and thus the point ofinterception can be estimated, for example from knowledge of the type oftarget.

An improved performance can be obtained in that the method and thearrangement can be used both in the trajectory phase of the missile andat least partially in its final phase.

The present invention will be described in more detail below withreference to the drawing, which shows an example of one advantageousembodiment.

FIG. 1 shows a scenario in which a missile is steering towards astationary target.

FIG. 2 shows a scenario in which the missile is steering towards amoving target in the form of an aircraft.

FIG. 3 shows an arrangement by means of which the missile guidance inFIGS. 1 and 2 is achieved.

In FIG. 1, 1 denotes a missile and 2 denotes a stationary target. Themissile 1 has information on its own position, velocity vector and thevelocity characteristic during the continuous flight of the missile 1.Access to these data is obtained through use of previously knownarrangements. For example, the missile, as in FIG. 3, has an inertialnavigation system 6, via which information is obtained on the positionand velocity vector of the missile. The characteristic for thecontinuous flight of the missile is obtained, for example, through amissile computer. In addition the missile has information on theposition of the target, for example as in FIG. 3 via a homing device orcommunications link 7. The homing device may be an IR sensor or radar,for example.

The missile 1 on the missile trajectory is designed to continuouslyhandle information on the missile 1 and the target 2, in order to guidethe missile towards the target, which is situated at point A in FIG. 1.As shown in FIG. 3, the missile in one example has a read-only memory11, in which software is stored, together with a processor 12, which isdesigned to carry out the instructions written into the software. Thesoftware is designed to read information on the missile and the targetinto a read-write memory 13 and from this information to calculate aflight time (ttg), which indicates the time it will take for the missileto travel to the target 2 by the shortest path. More precisely theflight time is obtained by resolving ttg from the following integral:∫_(t1 = 0)^(t2 = ttg)v̂(t)ts

where {circumflex over (ν)}(t) is the missile's further velocitycharacteristic and s is the distance between the missile and the targetat t=0, where t=0 is the current position on the missile trajectory.

The software is furthermore designed to calculate a fictitious point,marked B in FIG. 1, which is situated at a higher altitude than point Aand the distance of which to the latter is related to the calculatedflight time (ttg), the missile velocity vector being kept directedtowards the point B. In FIG. 1, Δ denotes the distance between point Aand point B. Point B is preferably situated along a vertical line thatpasses through the point A.

As stated previously, the distance Δ is related to the calculated flighttime (ttg). This relationship can be optimised with regard to therequired characteristics of the missile trajectory. For example, themissile trajectory can be optimised so that the missile maintains themaximum average velocity and/or so that it has the maximum finalvelocity. This optimisation is performed by simulations carried outbeforehand, in which account is taken of the characteristics andperformance of the missile, and external factors, for example. In thesimulations one or more parameters are developed for use by the missileby calculating the distance Δ between point A and point B. The distanceΔ may be described, for example, by a polynomial, in which

Δ(ttg)=p ₁*(ttg)² +p ₂*(ttg)+p ₃

and in which the parameters p₁, p₂ and p₃ are estimated by the saidsimulations on the basis of a selected optimisation criterion. Note thatthis is only an example. Δ as a function of the flight time ttg need notbe described by a polynomial. But Δ must decrease when the flight timedecreases to approach zero as ttg approaches zero.

The position of the point B is continuously updated during the missile'strajectory towards the target 2 based on updated calculations of theflight time ttg performed on the basis of updated information on theposition, velocity vector and further velocity characteristic of themissile. The missile velocity vector, which in an initial phase isdirected as shown by the dashed line 3, will, as ttg diminishes, bedirected closer and closer to point A. The dashed line 4 shows anexample of a missile trajectory towards point A. It is characteristic ofthe missile trajectory that directing the missile velocity vectortowards the fictitious point B, causes the missile to lob towards thetarget.

In FIG. 2, 1 denotes a missile and 2 denotes a target, as was the casein FIG. 1. Here, however, the target 2 is a moving target in the form ofan aircraft. As described earlier, the missile 1 has information on itsown position, velocity vector and the velocity characteristic during thecontinuous flight of the missile 1. In addition the missile continuouslyreceives information on the position and velocity vector of the target.As stated previously, the missile I has software, which is stored in aread-only memory 11, a processor 12 designed to carry out theinstructions written into the software and a read-write memory 13, intowhich information on the missile and the target is read. The software isdesigned, on the basis of the information read in, to predict a point ofinterception, marked A in FIG. 2, at which the trajectories of themissile 1 and the target 2 are simultaneously calculated to intersectand the missile is consequently expected to strike the target.

In this prediction use is made, on the one hand, of an assumptionregarding future behaviour of the target in order to estimate thetrajectory of the target, and on the other of a calculation of wherealong the estimated trajectory of the target the missile is calculatedto strike on the basis of the missile's information on its own position,velocity and further velocity characteristic. The assumption regardingthe future behaviour of the target can be made on the basis of a numberof different factors. If the target is stationary, (as in FIG. 1) thesolution becomes trivial, the target is assumed to be moving at avelocity 0. If the target is mobile it can be assumed to be moving witha constant velocity and direction, or with a constant velocity andconstant radius of curvature, or the assumption can also be based on thepremise that the target is being flown along the trajectory which makesit hardest for the missile to hit. In the latter case the missile islargely made to move towards the target in such a way that it will notmiss the target, even if the target is manoeuvred so that its trajectorymakes it as difficult as possible to hit.

When an assumption is made on future behaviour of the target, that is tosay when the trajectory in which the target is assumed to continue hasbeen determined, the point of interception A is predicted on the basisof this assumption of the future target behaviour. The prediction ismade by an iterative process in order to find a point along thetrajectory of the target, which the target and the missile can reachsimultaneously, that is to say a point to which a flight time for thetarget ttg_(target) is equal in length to a flight time for the missilettg_(missile).

A starting value ttg_(target″start) is first assigned to ttg_(target′)following which it is calculated at what point along its trajectory thetarget will be at this point in time. The starting valuettg_(target, start) may, for example, be stored beforehand in theread-only memory 16. It is then calculated what length of time themissile would need in order to reach the same point. We then call thistime ttg_(missile, start). If ttg_(target, start) andttg_(missile, start) do not coincide, a new value is calculated forttg_(target), which we then call ttg_(target, start+1). For example,ttg_(target, start+1) is calculated as the mean value ofttg_(target, start) and ttg_(missile,start). The process above is thenrepeated n number of times until ttg_(target, start+n) andttg_(missile, start+n) coincide or are sufficiently close to oneanother. We then have a value for the flight time ttg (i.e.ttg=ttg_(target)=ttg_(missile)), described in connection with FIG. 1,and the point at which the target and the missile are calculated to beat this time is the predicted point of interception A.

In addition the missile, as described earlier, is designed by means ofthe software to calculate a fictitious point of interception, denoted byB in FIG. 2, which is situated at a higher altitude than the predictedpoint of interception A and the distance Δ of which to that point isrelated to the calculated flight time (ttg). As previously, thisindicates the time it will take the missile to travel to the predictedpoint of interception A by the shortest path and, where the target is amoving target, is obtained from the iterative process in order topredict the point of interception A, as described.

The velocity vector of the missile is kept directed towards thefictitious point of interception B. This is continuously updated duringthe missile's trajectory towards the target 2, based on the updatediterative calculations of the flight time (ttg) performed on the basisof the updated information on the missile and the target. In oneembodiment the predicted point of interception A is updated once persecond and the fictitious point of interception is updated morefrequently, for example ten times per second. In this embodimentinvolving updates of the fictitious point of interception B in which thepoint of interception A is also updated, the flight time (ttg) isobtained when calculating the point of interception A, whereas in thoseupdates of B in which A is not updated the target 2 is regarded as astationary target, as described in connection with FIG. 1, the flighttime (ttg) being calculated according to the description of the flighttime calculation in connection with this figure.

In FIG. 2, 5 denotes an example of the target trajectory over a periodup to the time when the missile strikes the target at the point ofinterception A. As stated previously, the predicted point ofinterception A should be updated as the information on the target and onthe missile is updated and it has therefore shifted its position duringthe time up to impact.

FIG. 3 illustrates an arrangement 16, which has the elements that areneeded for performing the examples of a method according to the presentinvention described above. As stated previously, the reference numbers 6denote an inertial navigation system in the missile 1 and the referencenumbers 7 denote a homing device in the missile 1 and/or acommunications link. In one embodiment the missile's homing device 7 hasa relatively short range, for which reason it can only be used in finalphase guidance. When the missile is in the trajectory phase, it insteadreceives information on the target 2 via the communications link, forexample with a firing aircraft. The said aircraft has a long-range radarand can therefore supply information on the target when the missile isin the trajectory phase. Then when the missile is so close to the targetthat its own radar will function, the aircraft can in this embodimentseparate itself from the missile.

Furthermore there are interfaces 8 and 9 by means of which theinformation from the inertial navigation system 6 and the missilecomputer (not shown) and the homing device/the link 7 respectively canbe read into the read-write memory 13 via the processor 12. The softwarein the read-only memory 11 contains instructions for predicting a pointof interception at which the missile is expected to strike the targetusing the processor 12 on the basis of the information on the missileand the target previously referred to. In addition there areinstructions for calculating the flight time. In the flight timecalculation the distance s is calculated as the distance between themissile and the predicted point of interception at t=0, that is thecurrent situation of the missile on its trajectory.

The software also has instructions for calculating the fictitious pointof interception on the basis of the flight time obtained. In calculatingthe fictitious point of interception use is made of the parameters p₁,p₂, p₃ previously described, which are, for example, stored as constantsin the program. The read-only memory 11, the processor 12 and theread-write memory 13 are located in a computer 10.

The arrangement 16 furthermore has an interface 14, via which thefictitious point of interception calculated in the computer program isconverted to a form that an aiming device 15, connected to the interface14, can use in order to direct the missile velocity vector towards thesaid fictitious point.

As stated previously, the system described here has applications in thetrajectory phase and throughout or in part of the final phase.

What is claimed is:
 1. Method for guiding a missile (1) towards a target(2), the missile having information on its own position, velocity vectorand further velocity characteristic and continuously receivinginformation on the position and velocity vector of the target,characterised in that from the information which the missile has, apoint of interception (A) is predicted at which the missile is expectedto strike the target, a flight time is calculated which indicates thetime it will take for the missile to travel to the predicted point ofinterception (A), from a given criterion for the trajectorycharacteristics of the missile a fictitious point of interception (B) iscalculated, which is situated at a higher altitude than the predictedpoint of interception (A) and the distance (Δ) of which to the latterpoint is related to the calculated flight time and the missile is guidedtowards the said fictitious point.
 2. Method according to claim 1,characterised in that the fictitious point of interception is situatedalong a vertical line that passes through the predicted point ofinterception.
 3. Method according to claim 1, characterised in that theoptimisation criterion is maximum final velocity.
 4. An arrangement (16)in a missile (1) for guiding the latter towards a target (2), themissile having devices (6) designed to supply information on themissile's own position, velocity vector and further velocitycharacteristic, and devices (7) designed to continuously receiveinformation on the position and velocity vector of the target,characterised in that the arrangement (16) has first calculating devices(11, 12) designed, on the basis of the information which the missile (1)has, to predict a point of interception (A) at which the missile isexpected to strike the target, second calculating devices (11, 12)designed to calculate a flight time which indicates the time it willtake for the missile to travel to the predicted point of interception(A), third calculating devices (11, 12) designed, on the basis of agiven criterion for the trajectory characteristics of the missile, tocalculate a fictitious point of interception (B) which is situated at ahigher altitude than the predicted point of interception and thedistance of which to that point is related to the calculated flight timeand devices (15) designed to guide the missile towards the saidfictitious point.
 5. An arrangement according to claim 4, characterisedin that the fictitious point of interception (B) is situated along avertical line which passes through the predicted point of interception.6. An arrangement according to claim 4, characterised in that thereceiving device (7) is a homing device.
 7. An arrangement according toclaim 4, characterised in that the optimisation criterion is maximumfinal velocity.